With reference to FIG. 1, a ducted fan gas turbine engine generally indicated at 10 has a principal and rotational axis X-X. The engine comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, and intermediate pressure turbine 17, a low-pressure turbine 18 and an exhaust nozzle 19. A nacelle 21 generally surrounds the engine 10 and defines the intake 11, a bypass duct 22 and a bypass exhaust nozzle 23. A row of outlet guide vanes 24 is positioned in the bypass duct 22 rearward of the fan 12. A case 25 at the outer wall of the bypass duct 22 surrounds the fan 12 and the outlet guide vanes 24. The case 25 may be formed as different sections, and is strengthened to contain a fan blade in the unlikely event of a fan blade-off. The intermediate pressure compressor 13, high-pressure compressor 14, combustion equipment 15, high-pressure turbine 16, intermediate pressure turbine 17, low-pressure turbine 18 and exhaust nozzle 19 form the core engine 26.
The gas turbine engine 10 works in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the intermediate pressure compressor 13 and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 13 compresses the air flow A directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
The compressed air exhausted from the high-pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines 16, 17, 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines respectively drive the high and intermediate pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.
FIG. 2 shows schematically a perspective view from the rear of an engine similar to that shown in FIG. 1, but without the nacelle 21. The engine can be attached to an aircraft at an under wing pylon 27. Conventionally, the mounting assembly for the engine uses engine mounts which provide a detachable interface for the entire engine.
A forward engine mount 28 is attached to the case 25 and restrains the engine in side and vertical DOF (degrees of freedom). The case 25 and outlet guide vanes 24 can form a fan structure which is a major part of the engine architecture, the case 25, in particular, providing a mount ring for the forward engine mount 28, a rear fan case section (including a stiffener), and a forward containment fan case section. The fan structure can also include A frames 29 extending between the rear of the case 25 and the core engine 26. The fan structure provides a hard point for the front mount 28, and connects the case 25 to the core engine 26 in six DOE The core engine 26 is also attached to the pylon 27 at a tail bearing housing via a rear engine mount 30, this provides load transfer capability at the rear, and restrains the engine in side, vertical and roll DOFs. The axial DOF is restrained using thrust struts 31. The thrust struts 31 are attached to the rear engine mount 30 via a balance beam and extend forward to positions adjacent the A-frame 29 attachment positions on the core engine 26 to provide thrust load transfer capability only. The mounting assembly is effective in balancing flight generated loads (intake couple), with engine generated loads (thrust) to reduce “core bending”. Such bending can result in reductions in blade tip clearances, and is therefore detrimental to engine efficiency and performance as greater tip clearances are required to avoid rubs and tip wear.
As the split line between the aircraft and engine is at the engine mounts 28, 30, the assembly imposes a method of engine overhaul in which for major operations the whole engine is removed from under wing and transported to an overhaul base for maintenance work to be carried out. This can be both costly and time-consuming. In particular, as bypass ratios and fan diameters increase to meet growing demands in efficiency and noise reduction, it becomes a greater challenge to transport these large structures using both road and air freight.
The intake 11 is attached to the front of the case 25 such that normal aerodynamic loads and exceptional loads, e.g. due to fan blade off events, acting on the intake are transmitted from the intake to the case. However, as bypass ratios increase, the engine core diameter is reduced, and this reduction in core size has a negative effect on the structural ability of the engine to resist core bending. In particular, aerodynamic manoeuvring loads acting on the intake (caused, for example, by the aircraft angle of attack at takeoff) and transmitted to the case can lead to core bending. In addition, exceptional loads, such as gust loads, heavy landing loads, and fan blade off loads, can also act at times on the intake.
A further problem with the fan structure discussed above in relation to FIG. 2 is that the extra length of the case 25 to accommodate the A frames 29 can increase the length of the nacelle 21 and thereby reduce performance by increasing weight and drag. In addition, the A frames 29 cut across the air flow B through the bypass duct 22, and therefore impose an inherent drag penalty.
EP A 2202153 proposes a monolithic structure for mounting an engine to an aircraft.